1. Field of the Invention
The invention relates to a turbine blade and in particular, a blade for a gas turbine. The present invention also relates to a turbine comprising such a blade and an aircraft engine and in particular, a turbofan aircraft engine comprising a corresponding turbine. The present invention also relates to a method of reducing or eliminating the gap between a seal disposed on a flow-limiting wall of a turbine and a tip of a rotor blade or an outer shroud arranged on the rotor blade tip when the blade is at a temperature which is lower than the maximum operating temperature of the blade.
2. Discussion of Background Information
The materials currently employed for the blades of the first stage of a low pressure turbine are disadvantageous in that in the case of blades rotating at high speed or with high gas loads in the case of stationary vanes the creep resistance and tensile strength of these materials above about 1100° C. is insufficient. On the other hand, in order to further increase the efficiency of the currently available low pressure turbines (e.g., for high-speed turbofan aircraft engines) higher operating temperatures as well as higher rotating speeds are required. This causes a further increase in the thermal and mechanical stress of the components, in particular of the rotating blades and the stationary vanes of the first stages of the turbine. To reduce or prevent this increase in stress the components would have to be cooled, resulting in a decrease of efficiency and defeating the original purpose of increasing the efficiency.
Additionally, sealing systems in turbine components are to keep a gap between a rotating blade arrangement and a housing to a minimum and therefore are to guarantee a stable operation with a high degree of efficiency. Customarily, the rotating components of the turbine have sealing fins or sealing tips which, as is known, graze against or run in against seals disposed on the turbine wall, often honeycomb-shaped seals. For example, the rotating blades of the first stage of currently available low pressure turbines are usually made of materials such as Ni-based alloys which exhibit a thermal expansion coefficient of from 10×10−6 to 18×10−6 1/K in the temperature range from 20 to 1200° C. During operation at high temperatures (e.g., at about 1100° C. or higher) the blades expand and graze against or run in against the seals disposed on the flow-limiting wall of the turbine. This prevents an undesirable pressure loss in the corresponding turbine stage. However, at lower temperatures the blades contract again, resulting in a gap between the seal and the blade tip and thus, a pressure drop. The low pressure turbine can thus, not completely convert the energy present in the pressure gradient into work, thereby causing a loss of efficiency.
In view of the foregoing, it would be desirable to have available a turbine blade and a turbomachine such as a low pressure gas turbine for an aircraft engine which remedies the problems set forth above.